FR2789406A1 - ALCuMg ALLOY PRODUCT FOR AIRCRAFT STRUCTURE ELEMENT - Google Patents

Abstract

The subject of the invention is a rolled, extruded or forged product of AlCuMg alloy, treated by dissolving, quenching and cold pulling, intended for the manufacture of aircraft structural elements, of composition (% by weight) Fe <0.15 Si <0.15 Cu: 3.9 - 4.4 Mg: 1 - 1.5Mn 0.5 - 0.8 Zr: 0.08 - 0.15 other elements <0.05 each and < 0.15 in total, remains aluminum, and exhibiting a ratio Rm (L) / R0.2 (L)> 1.25. The invention applies more particularly to the manufacture of the lower surface of wings, exhibiting a set of properties (toughness, crack propagation speed, fatigue resistance, level of residual stresses) which are improved compared to alloy 2024. .

Description

I

AlCuMg alloy product for aircraft structural element.

TECHNICAL FIELD The invention relates to rolled, extruded or forged products of hardened and pulled AlCuMg alloy, intended for the manufacture of aircraft structural elements, in particular skin panels and airfoil pressure stiffeners, and having, compared to the products of the prior art used for the same application, an improved compromise between the properties of mechanical strength, formability, toughness, tolerance to damage and residual stresses. The designation of alloys and metallurgical states corresponds to the nomenclature of Aluminum

Association, taken over by European standards EN 515 and EN 573.

STATE OF THE ART The wings of large-capacity commercial aircraft include an upper part (or extrados) consisting of a skin made from thick sheets of alloy 7150 in the T651 state, or of alloy 7055 in the T7751 state. or 7449 in the T7951 state, and stiffeners made from profiles of the same alloy, and a lower part (or intrados) made up of a skin made from thick sheets of alloy 2024 in the T351 or 2324 state in state T39, and stiffeners made from sections of the

same alloy. The two parts are assembled by side members and ribs.

Alloy 2024 according to the designation of the Aluminum Association or the standard EN 573-3 has the following chemical composition (% by weight): Si <0.5 Fe <0.5 Cu: 3.8 - 4.9 Mg : 1.2 - 1.8 Mn: 0.3 - 0.9 Cr <0.10 Zn <0.25 Ti <0.15 Different variants have been developed and deposited with the Aluminum Association under the names 2224, 2324 and 2424, with in particular more limited silicon and iron contents. Alloy 2324 in the T39 state was the subject of Boeing patent EP 0038605 (= US 4,294,625), in which the improvement in the yield strength is obtained by work hardening using a pass of cold rolling after quenching. This work hardening tends to decrease the toughness and, to compensate for the decrease in toughness, the Fe, Si, Cu and Mg contents are reduced. Boeing has also developed alloy 2034 of composition: Si <0.10 Fe <0.12 Cu: -4.2-4.8 Mg: 1.3-1.9 Mn: 0.8- 1.3 Cr <0.05 Zn <0.20 Ti <0.15 Zr: 0.08-0.15 This alloy was the subject of patent EP 0031605 (= US 4336075). It has, compared to 2024 in the T351 state, a better specific elastic limit due to the increase in the manganese content and the addition of another anti-crystallizer.

o0 (Zr), as well as improved toughness and fatigue resistance.

The patent EP 0473122 (= US 5213639) of Alcoa describes an alloy, registered with the Aluminum Association as 2524, of composition: Si <0.10 Fe <0.12 Cu: 3.8 - 4.5 Mg: 1 , 2 - 1.8 Mn: 0.3 - 0.9 which may optionally contain another anti-crystallizer (Zr, V, Hf, Cr, Ag or Sc). This alloy is intended more particularly for thin sheets for fuselage and exhibits toughness

and improved crack propagation resistance compared to 2024.

The applicant's patent application EP 0731185 relates to an alloy, subsequently registered under No. 2024A, of composition: Si <0.25 Fe <0.25 Cu: 3.5 - 5 Mg: 1 - 2 Mn <0.55 with the relationship: 0 <Mn - 2Fe <0.2 The thick sheets of this alloy exhibit both improved toughness and

reduced level of residual stresses, without loss on other properties.

Alcoa's US Pat. Nos. 5863359 and US 5865914 respectively relate to an aircraft wing comprising a lower surface made of an alloy of composition Cu: 3.6 - 4 Mg: 1 - 1.6 (pref: 1.15 - 1.5) Mn : 0.3 - 0.7 (pref .: 0.5 - 0.6) Zr: 0.05 0.25 and preferably Fe <0.07 and Si <0.05 exhibiting both the following properties - Ro , 2 (LT)> 60 ksi (414 MPa) and Kt1 (LT)> 38 ksiVinch (42 MPa'Im), and a method of manufacturing an inner surface element having a Ro, 2 (LT)> 60 ksi comprising the casting of an alloy of the preceding composition, a homogenization between 471 and 482 C, a hot transformation at a temperature> 399 C, a mraise in solution above 488 C, a quenching, a cold hardening preferably of

more than 9% and a traction of at least 1%.

Problem posed For the construction of new large-capacity commercial airplanes, it is of course imperative to limit the weight, so that the specifications of the manufacturers impose higher typical stresses for the wing panels, which results in lower minimum values. high for the static mechanical characteristics and damage tolerance of the aluminum alloy products used. The use of products hardened in the T39 state, such as those recommended in US Pat. Nos. 5863359 and US 5865914, although it leads to high elasticity limits Ro, 2, however, presents a certain number of drawbacks for other important usage properties in the intended application. Indeed, this results in a plastic deviation, that is to say a difference between the tensile strength R ,, and the elastic limit R0.2, very reduced, which results in a lower cold formability. and poorer resistance to fatigue crack propagation with variable amplitude loading. Indeed, the slowing down of the propagation of cracks after

partial overload is less if the plastic gap is reduced.

In addition, larger parts must be machined without distortion in thicker sheets, which implies better control of the stress level.

residuals. However, the T39 state has proved unfavorable from this point of view.

The aim of the invention is therefore to provide products made of AICuMg alloy in the quenched and cold-deformed state, intended for the manufacture of the intrados of aircraft wings, and exhibiting, compared with similar products of the prior art, a more favorable compromise for all the use properties: mechanical strength, crack propagation speed, toughness, fatigue resistance, and residual stress rate. OBJECT OF THE INVENTION The subject of the invention is a rolled, extruded or forged product made of AICuMg alloy, treated by dissolving, quenching and cold pulling, intended for the manufacture of aircraft structural elements, of composition ( % by weight): Fe <0.15 Si <0.15 Cu: 3.9-4.4 (pref: 4.0-4.3) Mg: 1.0-1.5 Mn: 0.5 - 0.8 Zr: 0.08 - 0.15 other elements: <0.05 each and <0.15 in total, exhibiting an RfI (L) / Roi, 2 (L) ratio of tensile strength in the direction L to

the elastic limit in the L direction, greater than 1.25 (and preferably 1.30).

It also relates to a rolled product (a sheet) of the same composition with a thickness of between 6 and 60 mm and having in the quenched and pulled state at least one of the following groups of properties; a) Tensile strength R, (L)> 475 Mpa and yield strength RO, 2 (L)> 370 MPa b) Plastic deviation Rn - Ro, 2 directions L and TL> 100 MPa c) Intensity factor critical (LT direction) Kó> 170 MPa'Jm and Ko> 120 MPaVm

(measured according to standard ASTM E561 on notched specimens taken at quarter-

thickness with parameters B = 5 mm, W = 500 and 2ao = 165 mm) d) Crack propagation speed (LT) da / dn, measured according to standard ASTM E 647 on notched specimens taken at quarter thickness with W = 200 mm and B = 5mm: <10-4 mm / cycle for AK = 10 MPa'm <2.5 104 mm / cycle for AK = 15 MPaVm and <5 104 mm / cycle for AK = 20 MPaIm This sheet presents also a level of residual stresses such as the deflection f measured in the L and TL directions after mid-thickness machining of a bar resting on two supports spaced apart by a length I is such that: f <(0.14 12) / ef being measured in microns, the thickness e of the sheet and the length I

being expressed in mm.

A subject of the invention is also a process for manufacturing a rolled, extruded or forged product comprising the following steps: - casting of a plate or a billet of the indicated composition, - homogenization of this plate or billet between 450 and 500 C, - hot and possibly cold transformation to the desired product, - solution at a temperature between 480 and 505 C, - cold water quenching, cold pulling with at least 1.5% permanent deformation,

- natural aging at room temperature.

Description of the invention

The chemical composition of the product differs from that of the usual 2024 by a reduced content of iron and silicon, with preferably Fe + Si <0.15%, a higher content of manganese and an addition of zirconium. Compared to 2034, we have a lower manganese content and a slightly reduced copper content. Compared to the composition of the alloys described in US Pat. No. 5,863,359 and US Pat. No. 5,865,914, the copper content is higher, which makes it possible to compensate, for mechanical strength, for the lower cold work hardening after quenching. Surprisingly, this narrow range of composition (in particular with regard to manganese), associated with modifications to the production range, leads, compared to the prior art, to a significant improvement in the compromise between strength. mechanical, elongation and damage tolerance under the operating conditions of a large capacity civil aircraft. In addition, and quite unexpectedly, a low rate of residual stresses is observed for thick products, allowing distortion-free machining of large parts. The manufacturing process comprises the casting of plates, in the case where the product to be manufactured is a rolled sheet, or of billets in the case of a extruded profile or a forged part. The plate or the billet is scalped, then homogenized between 450 and 500 C. The hot transformation is then carried out by rolling, extruding or forging. This transformation is preferably carried out at a temperature higher than the temperatures usually used, the outlet temperature being greater than 420 C and preferably at 440 C so as to obtain on the treated product a slightly recrystallized structure, with a recrystallization rate. a quarter thickness less than 20%, and preferably 10%. The rolled, spun or forged semi-finished product is then put into solution between 480 and 505 C, so that this dissolution is as complete as possible, that is to say that the maximum of potentially soluble phases, in particular the precipitates A12Cu and AI2CuMg, ie effectively in solid solution. The quality of the solution can be assessed by differential enthalpy analysis (DEA) by measuring the specific energy using the area of the peak on the

thermogram. This specific energy should preferably be less than 2 J / g.

Then one proceeds to quenching in cold water, and a controlled traction leading to a permanent elongation of at least 1.5%. The product finally undergoes natural aging at room temperature. The products according to the invention exhibit markedly improved static mechanical characteristics compared to the 2024-T351 alloy, currently used for airplane wing lower surfaces, and barely weaker than those of 2034-T351. The high plastic gap and elongation of the material results in excellent cold formability. The toughness, measured by the critical stress intensity factors in plane stress Kc and Kco is more than 10% higher than that of 2024 and 2034, and the crack propagation speed da / dn is markedly improved compared to these two alloys, in particular for the high values of AK, and for loadings with variable amplitude. The fatigue lifetimes, measured on notched specimens taken at mid-thickness in the L direction, are also improved by more than 20% compared to 2024 and 2034. Finally, the level of residual stresses, measured by the arrow f after machining at mid-thickness of a bar resting on two distant supports of a length 1, is rather low, whereas one could have expected on the contrary with a fiber structure. This deflection, measured in microns, is always less than the quotient (0.14 12) / e, the length 1 and

the thickness e of the sheet being expressed in mm.

All of these properties mean that the products according to the invention are particularly well suited to the manufacture of aircraft structural elements, in particular the lower surfaces of the wings, but also sections for wing box, for sole of spars. and assembled ribs and fuselage skins and stiffeners.

Examples

Three plates 1450 mm wide and 446 mm thick, respectively, made of alloy 2024, 2034 and alloy according to the invention were cast. The chemical compositions (% by weight) of the alloys are given in Table 1:

Table 1

Si Fe Cu Mg Mn Zr alloy

2024 0.12 0.20 4.06 1.36 0.54 0.002

2034 0.05 0.07 4.30 1.34 0.98 0.14

invention 0.06 0.08 14.14 1.26 0.65 0, 102 The plates were scalped, then homogenized under the following conditions For 2024, 2 h at 495 C then 5 h at 460 C For 2034, 5 h at 497 C For the alloy according to the invention, rise in 12 h and hold for 6 h at 483 C Part of the sheets was then hot rolled to a thickness of 40 mm by successive passes of the order of 20 mm. Another part of the sheets was hot rolled up to 15 mm. For the alloy according to the invention, the hot-rolling inlet temperature was 467 C, the outlet temperature at 40 mm was 465 C and that at

mm 444 C.

The sheets were dissolved under the following conditions 3h and 6h at 497 C for the 2024 sheets with a respective thickness of 15 and 40 mm, 2h and 5h at 499 C for the 2034 sheets with a thickness of 15 and 40 mm

9 am at 497 C for the sheets according to the invention.

After quenching in cold water, all the sheets were then subjected to a controlled traction at

2% permanent elongation.

The static mechanical characteristics in the L and TL directions were measured on the sheets, namely the tensile strength RP (in MPa), the conventional yield strength at 0.2% Ro, 2 (in MPa) and l elongation at break A (in%). The results are collated in Table 2

Table 2

Alloy Thickness Sens. Ro, 2 A

2024 40 L 468 362 20.0

2024 40 TL 469 330 17.4

2024 15 L 462 360 21.2

2024 15 TL 467 325 17.6

2034 40 L 534 416 11.2

2034 40 TL 529 393 12.0

2034 15 L 548 431 13.8

2034 15 TL 531 395 14.6

Invention 40 L 510 384 15, 4 Invention 40 TL 475 336 18.9 Invention 15 L 501 390 16.7 Invention 15 TL 491 351 19.1 The toughness was also measured by the critical intensity factors in plane stress Kc and Ko (in MPalm) in the LT direction, according to standard ASTM E 561, on CCT specimens, taken at quarter-thickness, width W = 500 mm, thickness B = 5 mm, and a central notch machined by spark erosion 2ao = 165 mm, enlarged by fatigue test up to 170 mm. The results are given in Table 3 Table 3 Alloy Thickness K Kzo

2024 40 143.4 105.2

2034 40 128.8 97.8

Invention 40,179.7 122

2034 15 136.4 103.7

Invention 15 173.6 124.3 The fatigue crack propagation speed da / dn in the LT direction (in mm / cycle) was also measured for different values of AK (in MPaJm) according to the ASTM E 647 standard. uses for this 2 CCT specimens of width W = 200 mm and thickness B = 5 mm, taken at quarter sheet thickness in the LT direction. The length of the central notch machined by spark erosion is 30 mm, and this notch is enlarged by fatigue test to 40 mm. The cracking rate measurement test is carried out on an MTS machine with a stress in R = 0.05 and a stress of 40 MPa, calculated to obtain an AK value of 10 MPa'Im for the length

40 mm starting notch (results in Table 4).

Table 4

Alloy Ep. AK = 10 AK = 12 AK = 15 AK = 20 AK = 25 2024 40 9 I0o5 1.5 10-4 3.0 10-4 6 10-4 9 10o-3

2034 40 8 10-5 1.5 104 3 104 5.7 10-4 1.7 10-3

Inv. 40 5.5 10-5 1.7 10-4 2.0 10-4 4.0 104 7.8 104 2034 15 8 i0o- 1.5 10-4 3 104 5.2 10-4 2.1 03 Inv. 15 4.9 10 '6.0 10'5 1.3 10'4 2.5 10-4 5.4 10-4 Fatigue tests according to the Airbus AITM 1-0011 specification were carried out on test specimens with holes 230 mm long, 50 mm wide and 7.94 thick

mm, taken at mid-thickness of the L direction plate. The diameter of the hole is 7.94 mm.

An average full specimen stress of 80 MPa was applied with 4 levels of alternating stresses: 85 MPa, 55 MPa, 45 MPa and 35 MPa for sheets of 40

mm, 110, 85, 55 and 45 MPa for 15 mm sheets, with 2 specimens per level.

The average lifetime values (in number of cycles) are shown in Table 5. It is noted that, for specimens with a notch factor K. = 2.5, the duration

fatigue life is improved by more than 20% compared to alloy 2024.

Table 5

alloy Thickness 80 + 85 80 + 55 80 + 45 80 + 35 mm MPa MPa MPa MPa

2024 40 36044 159721

2034 40 30640 125565 340126 839340

invention 40 42933 219753 392680 1018240

2034 15 41040 204038 352957

invention 15 45841 241932 429895 Finally, the arrows f in the direction L and TL were measured, as well as the rate of recrystallization (in%) at the surface, at quarter-thickness and at mid-thickness, determined by

image analysis after chemical attack of the sample.

The deflection f is measured as follows. Two bars are taken from the sheet of thickness e, one called a bar L direction, of length b in the direction of the length of the sheet (direction L), of width 25 mm in the direction of the width of the sheet (direction TL) and thickness e according to the full thickness of the sheet (direction TC), the other, called bar

TL direction, having 25 mm in the L direction, b in the TL direction and e in the TC direction.

Each bar is machined to mid-thickness and the deflection is measured at mid-length of the bar. This deflection is representative of the level of internal stresses of the sheet and of its ability not to deform during machining. The distance I between the supports was 180 mm and the length b of the bars was 200 mm. Machining is machining

progressive mechanics with passes of about 2 mm. The measurement of the deflection at mid

length is performed using a comparator with a resolution of one micron. The results concerning the arrows and the recrystallization rates are given in Table 6. It

Table 6

alloy Thickness fL (inm) fTL (km) Recr. Recruit rate Recruit rate (Surf.)% (1/4 ep.)% ('/ 2 ep.)%

2024 40 210 120 79 58 30

2034 40 147 129 12 0 0

Invention 40 86 75 46 5 2

Claims (10)

Claims i. Rolled, extruded or forged product in AICuMg alloy, treated by dissolving, quenching and cold pulling, intended for the manufacture of aircraft structural elements, of composition (% by weight): Fe <0.15 Si < 0.15 Cu 3.9-4.4 (pref 4.0 - 4.3) Mg: 1-1.5 Mn: 0.5 - 0.8 Zr: 0.08 - 0.15 other elements: < 0.05 each and <0.15 in total, and exhibiting an R ,, (L) / R0.2 (L) ratio> 1.25 (preferably> 1.30). 2. Product according to claim 1, characterized in that Fe + Si <0.15%. 3. Rolled product with a thickness of 6 to 60 mm according to one of claims I or 2, having in the hardened and pulled state a tensile strength R, (L)> 475 MPa and a yield strength RO, 2 (L)> 370 MPa 4. Laminated product 6 to 60 mm thick according to one of claims I to 3, exhibiting in the hardened and pulled state a plastic gap between the resistance to rupture R, and the yield strength RO2 in the L and TL directions> 100 MPa. 5. Laminated product 6 to 60 mm thick according to one of claims 1 to 4, exhibiting in the quenched and pulled state a critical intensity factor (L-T) K> MPam and KIo> 120 MPaVm, measured according to standard ASTM E 561 on notched specimens taken at quarter-thickness with the parameters W = 500 mm, B = 5 mm and 2ao = 165 mm. 6. Laminated product according to one of claims 1 to 5, characterized in that it has a critical intensity factor (L-T direction) Kc or Ko increased by at least 10% compared to alloy 2024 under the same conditions. 7. Laminated product 6 to 60 mm thick according to one of claims I to 6, exhibiting in the hardened and pulled state a crack propagation speed (L- T) da / dn, measured according to standard ASTM E 647 on notched specimens taken at quarter-thickness with W = 200 mm and B = 5mm: <10-4 mm / cycle for AK = 10 MPaVm <2.5 10- 4 mm / cycle for AK = 15 MPaIm and <5 104 mm / cycle for AK = 20 MPal / m 8. Rolled product according to one of claims I to 7, characterized in that it has an arrow f measured in the direction L and TL after machining at mid-thickness of a bar resting on two distant supports with a length 1 less than (0.14 12) / e, f being measured in microns, the thickness e of the sheet and the length I being expressed in mm 9. Laminated product according to one of claims 3 to 8, having a lifespan average fatigue, measured on a notched specimen taken halfway thickness direction L, increased by more than 20% compared to alloy 2024. 10. A method of manufacturing a product according to one of claims I to 9, comprising the following steps: - casting of a plate of the composition indicated - homogenization of this plate between 450 and 500 ° C., hot transformation, and optionally cold, by rolling, extrusion or forging until the desired product is reached, solution at a temperature between 480 and 505 C, - quenching in cold water, - cold traction up to more than 1.5% permanent deformation, - natural aging at room temperature. 1 1. The method of claim 10, characterized in that the hot transformation is done with an outlet temperature> 420 C, and preferably> 440 C. 12. Use of sheets according to one of claims 3 to 9 for the manufacture of aircraft wing pressure panel skin. 13. Use of profiles according to one of claims I or 2 for the manufacture of Airplane wing or aircraft fuselage lower surface stiffeners.